Referring to FIG. 1, a gas turbine engine is generally indicated at 10 and comprises, in axial flow series, an air intake 11, a propulsive fan 12, an intermediate pressure compressor 13, a high pressure compressor 14, combustion equipment 15, a high pressure turbine 16, an intermediate pressure turbine 17, a low pressure turbine 18 and an exhaust nozzle 19.
The gas turbine engine 10 works in a conventional manner so that air entering the intake 11 is accelerated by the fan 12 which produce two air flows: a first air flow into the intermediate pressure compressor 13 and a second air flow which provides propulsive thrust. The intermediate pressure compressor compresses the air flow directed into it before delivering that air to the high pressure compressor 14 where further compression takes place.
The compressed air exhausted from the high pressure compressor 14 is directed into the combustion equipment 15 where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive, the high, intermediate and low pressure turbines 16, 17 and 18 before being exhausted through the nozzle 19 to provide additional propulsive thrust. The high, intermediate and low pressure turbine 16, 17 and 18 respectively drive the high and intermediate pressure compressors 14 and 13, and the fan 12 by suitable interconnecting shafts.
It will be appreciated from above that the turbine blades require appropriate mounting in order to allow rotation for operational performance in creating a propulsive axial gas flow, but also that the blades must be appropriately cooled. It will be understood that turbine engine efficiency is closely related to operational temperatures and that acceptable operational temperatures are dictated to a significant extent by the material properties of the components. In such circumstances by appropriate cooling it is possible to operate these components near to and occasionally exceeding the melting points for the materials from which they are constructed.
In order to provide cooling, generally coolant air is taken from the compressor stages of a gas turbine engine. Thus, this drainage of compressed coolant air reduces engine efficiency. It is an objective to utilise coolant air flows as effectively as possible in order to minimise the necessary coolant flow to achieve a desired level of component cooling for operational performance. In such circumstances generally there are relatively intricate coolant passageways provided within the engine components which are arranged to provide cooling as the coolant passes through these passages as well as provide generally nozzle projection of the coolant flows where required into cavities in order to create turbulence with hot gas flows for a diluted cooling effect.
FIG. 2 illustrates a schematic cross-section of a prior cooling arrangement as a schematic cross-section. Thus a blade root 1 forms a shank with a locking plate 2 presented across the root 3 of the blade. With a gas turbine engine, banks of turbine blades are provided and it is necessary to provide sealing between each turbine stage of the engine. Seals 4 are provided in the form of a labyrinth seal arrangement with coolant airflow in the direction of arrowhead 5 presented upwardly into the cavity 6 formed between the mounting disc 7 for the blade 1 and the bottom of a nozzle vane defining the turbine stages. As can be seen there is a gap 8 through which hot gas is ingested to the cavity 6. It is found that the coolant leakage flow 9 generally creates a barrier layer around the surfaces of the cavity 6 particularly on the rotating surface. Previously, the coolant air 5 has been arranged to prevent excessive hot gas ingestion 8.
The lock plate acts to secure location of the blade shank such that coolant flow is contained or at least restricted below the blade shank. It will be appreciated that as described in U.S. Pat. No. 6,290,464, an area 10 adjacent the lock plate is typically of what is known as a fir tree root nature and designed to allow coolant air to flow across it at its surface and possibly through passages (not shown) in the fir tree root in order to provide cooling. As turbine engines rotate about a central shaft they are inherently circumferential and it is therefore necessary that the lock plate is segmented. In such circumstances the gaps between the lock plates allow coolant leakage into the cavity. U.S. Pat. No. 6,290,464 describes provision of an outlet nozzle in order to project coolant flow through the fir tree root coolant passages into such a cavity in order to create turbulence and therefore cooling within that cavity. Such an approach does not utilize the boundary layer created by the lockplate leakage to protect the disc rim from ingestion of hot gas through the gap.